Geared gas turbine engine

ABSTRACT

The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

The present disclosure relates to a geared gas turbine engine for anaircraft.

Turbofan gas turbine engines for aircraft propulsion have many designfactors that affect the overall efficiency and power output or thrust.To enable a higher thrust at a high efficiency, a larger diameter fanmay be used. As the diameter of the fan is increased, however, therequired lower speed of the fan tends to conflict with the requirementsof the turbine component the core shaft is connected to, typically a lowpressure turbine. A more optimal combination can be achieved byincluding a gearbox between the fan and the core shaft, which allows thefan to operate at a reduced rotational speed at higher efficiency, andtherefore enables a larger size fan, while maintaining a high rotationalspeed for the low pressure turbine, enabling the overall diameter of theturbine to be reduced and a greater efficiency to be achieved with fewerstages.

A high propulsive efficiency for a geared gas turbine engine is achievedthrough a high mass flow through the engine. This may be enabled in partby increasing the bypass ratio of the engine, which is the ratio betweenthe mass flow rate of the bypass stream to the mass flow rate enteringthe engine core. To achieve a high bypass ratio with a larger fan whilemaintaining an optimum gearing ratio and fan speed, the size of theengine core, in particular the low pressure turbine, may need toincrease, which would make integration of a larger fan engine underneathan aircraft wing more difficult. A general problem to be addressedtherefore is how to achieve a high propulsive efficiency for a largergeared gas turbine engine while enabling the engine to be integratedwith an aircraft.

According to a first aspect there is provided a gas turbine engine foran aircraft comprising:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft to drive        the fan at a lower rotational speed than the core shaft, the        gearbox having a gear ratio of around 3.4 or higher,    -   wherein the gas turbine engine is configured such that a jet        velocity ratio between a first jet velocity through a bypass        duct of the engine and a second jet velocity through the engine        core is within a range from around 0.75 to around 0.82 at cruise        conditions.

An advantage of configuring the engine so that the jet velocity ratio iswithin the above range is that a high bypass ratio can be maintained,enabling the engine to be efficient with a low specific thrust, andmaintaining a high gearing ratio so that the size of the engine core canbe kept small. The rotational speed of the low pressure turbine of theengine core can therefore be kept high, enabling it to be kept smaller,which avoids problems with integration of the engine under an aircraftwing.

The gear ratio of the gearbox enables the fan to rotate more slowly thanthe low pressure turbine of the engine core. As the gear ratioincreases, the advantage of having a gearbox decreases due to theincreased losses in the gearbox and increased wear resulting fromcomponents such as planetary gears that need to be smaller and rotatefaster. The gear ratio may therefore advantageously be around 5.0 or 4.5or less, i.e. resulting in a range for the gear ratio of between around3.4 to around 5.0 or around 4.5.

To preserve propulsive efficiency, the kinetic energy difference betweenthe cold and hot stream jets, i.e. the jets from the bypass and coreexhaust, should be minimised. Decreasing the jet velocity ratio belowaround 0.75 will tend to reduce propulsive efficiency and therefore havean increased detrimental effect on fuel burn rate. The jet velocityratio may therefore be kept within a range from around 0.75 to around0.82.

According to a second aspect there is provided a gas turbine engine foran aircraft comprising:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft to drive        the fan at a lower rotational speed than the core shaft, the        gearbox having a gear ratio of between around 3.2 and around        3.8, optionally around 3.3 to around 3.8,    -   wherein the gas turbine engine is configured such that a jet        velocity ratio between a first jet velocity through a bypass        duct of the engine and a second jet velocity through the engine        core is within a range from around 0.75 to around 1.0 at cruise        conditions.

The jet velocity ratio, R_(J), may be defined as:

$R_{J} = \frac{V_{B}C_{B}}{V_{C}C_{C}\eta_{LPT}\eta_{F}}$

-   -   where V_(B) is the fully expanded first jet velocity, C_(B) is a        thrust coefficient of the bypass nozzle, V_(C) is the fully        expanded second jet velocity, C_(C) is a thrust coefficient of        the core exhaust nozzle, η_(LPT) is an isentropic efficiency of        a lowest pressure turbine of the engine core and η_(F) is an        isentropic efficiency of compression of air into the bypass duct        by the fan. The fully expanded jet velocity may be defined as        the axial jet velocity at the point where the exhaust jet has        expanded to ambient pressure. The term nozzle thrust coefficient        (C_(B) and C_(C)) as used herein has the standard meaning in the        art, as understood by the skilled person.

The gearbox may be an epicyclic gearbox comprising an input sun gearconnected to the core shaft, a plurality of planetary gears connected bya carrier arm and an outer annulus ring, the fan being connected to thecarrier arm. Alternatively, the gearbox may be an epicyclic gearboxcomprising an input sun gear connected to the core shaft, a plurality ofplanetary gears connected by a carrier arm and an outer annulus ring,the fan being connected to the outer annulus ring.

The fan may have an outer diameter of between around 320 cm and around400 cm. In some examples the fan may have an outer diameter of betweenaround 330 cm and around 370 cm.

In some examples, where the turbine is a first turbine, the compressor afirst compressor, and the core shaft a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a third aspect there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine comprising:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft to drive        the fan at a lower rotational speed than the core shaft, the        gearbox having a gear ratio of around 3.4 or higher,    -   wherein the method comprises operating the gas turbine engine to        provide propulsion under cruise conditions such that a jet        velocity ratio between a first jet velocity exiting from a        bypass duct of the engine and a second jet velocity exiting from        an exhaust nozzle of the engine core is within a range from        around 0.75 to around 0.82

According to a fourth aspect there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine comprising:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft to drive        the fan at a lower rotational speed than the core shaft, the        gearbox having a gear ratio of between around 3.2 and around        3.8, optionally around 3.3 to around 3.8,    -   wherein the method comprises operating the gas turbine engine to        provide propulsion under cruise conditions such that a jet        velocity ratio between a first jet velocity exiting from a        bypass duct of the engine and a second jet velocity exiting from        an exhaust nozzle of the engine core is within a range of from        around 0.75 to around 1.0.

The optional and advantageous features described above in relation tothe first and second aspects may be applied also to the method accordingto the third and fourth aspects.

Cruise conditions may be defined as a forward Mach number of between 0.7and 0.9 at an altitude of between 10000 m and 15000 m. Other conditionssuch as ambient temperature and pressure are largely dependent on thealtitude.

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example 3.2 or higher, for example in the range of from 3.2to 5.0, for example on the order of or at least 3.2, 3.3, 3.4, 3.5, 3.6,3.7, 3.8, 3.9, 4.0, 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9 and 5.0.The gear ratio may be, for example, between any two of the values in theprevious sentence.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches),350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380(around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm(around 160 inches) or 420 cm (around 165 inches). The fan diameter maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 0.28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 13 to 16, or 13 to 15, or 13 to 14. Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is an example plot of change in fuel burn as a function of jetvelocity ratio;

FIG. 5 is a schematic drawing of an aircraft having a gas turbine enginemounted thereon; and

FIG. 6 is a schematic drawing illustrating the concept of a fullyexpanded jet velocity.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is com busted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 illustrates an example plot of change in fuel burn, ΔFB, as afunction of jet velocity ratio, R_(J)., other factors being constant.The change in fuel burn contribution from propulsive efficiency 401 isdetermined relative to an optimum value for the jet velocity ratio ofaround 1.0, with an increase in fuel burn above and below this value.The change in fuel burn contribution from the IP turbine, represented byline 402 in FIG. 4 (for example relative to an optimum IP turbine, whichmay be achieved at an very low jet velocity ratio off the left hand sideof the graph) may be due to, for example, the change in weight and/orsize of the IP turbine as the velocity ratio changes, other factorsbeing constant. In this regard, an increased velocity ratio maygenerally result in a larger and/or heavier IP turbine, which may alsohave an adverse impact on the installation of the engine with theaircraft. Factors that may affect the jet velocity ratio include therelative rotational speeds of the fan and turbine and the areas of theexhaust nozzles for the bypass and core exhausts.

For a higher gearing ratio, i.e. around 3.2, 3.3, 3.4 and above, forexample up to around 3.8 or in some cases even higher, the jet velocityratio tends to be around 1.0 or below. As the jet velocity ratiodecreases below 1.0, the fuel burn contribution from propulsiveefficiency 401 increases, and at a higher rate than for the portionabove 1.0. To maintain this loss to within around 0.5% for such anarrangement, it can be seen from FIG. 4 that the jet velocity ratiowould need to be kept within around 0.8 to around 1.0, and for a jetvelocity ratio of around 0.75 and below, the fuel burn contribution frompropulsive efficiency becomes dominant, rising to around 0.7% and above.A lower limit for the jet velocity ratio of around 0.85 or 0.90 may beused to keep the fuel burn contribution from propulsive efficiency toaround 0.25% or below. It has been found that further decreasing the jetvelocity ratio enables a higher gear ratio to be used and/or decreasesthe pressure ratio across the IP turbine, thereby allowing for asmaller, faster and/or lighter IP turbine, reflected in a lowercontribution to fuel burn loss 402 by the IP turbine. Such an IP turbinemay have installation benefits when installed on an aircraft, forexample in terms of the ability to better optimize the position theengine relative to a wing. A range of around 0.75 to around 1.0 or 0.75to around 0.82 for the jet velocity ratio has been found to beadvantageous for engines with higher gear ratios as defined herein.

A lower gear ratio of the gearbox typically results in values for thejet velocity ratio of 1.0 or greater. For gearboxes having such ratios,to keep the fuel burn loss due to propulsive efficiency to within around0.5% or less of the optimum, it can be seen from FIG. 4 that the jetvelocity ratio should be between around 1.0 and around 1.3. As the jetvelocity ratio increases further, the increase in fuel burn contributionboth from the propulsive efficiency (line 401) and from the IP turbine(line 402) becomes greater. An upper limit for the jet velocity ratio isaround 1.2 keeps the increase in fuel burn due to propulsive efficiencyto around 0.25-0.3%.

For a given set of gears making up an epicyclic gearbox, a planetarydriving arrangement will produce a higher gearing ratio than a stardriving arrangement. A star arrangement may be generally preferred incombination with a jet velocity ratio of around 1.0 and above, and aplanetary arrangement may be generally preferred for a jet velocityratio of around 1.0 and below. However, it will be appreciated that starand planetary gearboxes may be used outside of these preferred ranges.

FIG. 5 illustrates an example aircraft 50 having a gas turbine engine 10attached to each wing 51 a, 51 b thereof. When the aircraft is flyingunder cruise conditions, as defined herein, each gas turbine engine 10operates such that a jet velocity ratio between a first jet velocityexiting from a bypass duct of the engine 10 and a second jet velocityexiting from an exhaust nozzle 20 of the engine core is within a rangefrom around 1.0 to around 1.3.

FIG. 6 illustrates an example exhaust nozzle 60 of a gas turbine engine.The pressure Pj at the exit or throat 61 of the exhaust nozzle 60 isgreater than the ambient pressure Pa around the engine. At some distanceaway from the nozzle exit 61 the jet pressure will equalise with theambient pressure, i.e. Pj=Pa. The fully expanded jet velocity is definedas the jet velocity 62 at this point, i.e. the jet velocity along theaxis of the engine at a minimum distance from the exhaust nozzle wherethe pressure is equal to ambient pressure.

It will be understood that the invention is not limited to theembodiments above- described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox thatreceives an input from the core shaft to drive the fan at a lowerrotational speed than the core shaft, the gearbox having a gear ratio ofaround 3.4 or higher, wherein the gas turbine engine is configured suchthat a jet velocity ratio between a first jet velocity through a bypassduct of the engine and a second jet velocity through the engine core iswithin a range from 0.75 to around 0.82 at cruise conditions.
 2. The gasturbine engine of claim 1, wherein the gear ratio is within a range fromaround 3.4 to around 5.0 or around 4.5.
 3. A gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and a gearbox that receives an input from the core shaft todrive the fan at a lower rotational speed than the core shaft, thegearbox having a gear ratio of between around 3.2 and around 3.8,wherein the gas turbine engine is configured such that a jet velocityratio between a first jet velocity through a bypass duct of the engineand a second jet velocity through the engine core is within a range from0.75 to around 1.0 at cruise conditions.
 4. The gas turbine engine ofclaim 1 wherein the jet velocity ratio, R_(J), is defined as:$R_{J} = \frac{V_{B}C_{B}}{V_{C}C_{C}\eta_{LPT}\eta_{F}}$ whereV_(B) is the fully expanded first jet velocity, C_(B) is a thrustcoefficient of the bypass nozzle, V_(C) is the fully expanded second jetvelocity, C_(C) is a thrust coefficient of the core exhaust nozzle,η_(LPT) is an isentropic efficiency of a lowest pressure turbine of theengine core and η_(F) is an isentropic efficiency of compression of airinto the bypass duct by the fan.
 5. The gas turbine engine of claim 1wherein the gearbox is an epicyclic gearbox comprising an input sun gearconnected to the core shaft, a plurality of planetary gears connected bya carrier arm and an outer annulus ring, the fan being connected to thecarrier arm.
 6. The gas turbine engine according to claim 1, wherein thefan has an outer diameter of between around 220 cm and around 390 cm. 7.The gas turbine engine according to claim 6, wherein the fan has anouter diameter of between around 240 cm and around 280 cm or betweenaround 330 cm and around 380 cm.
 8. The gas turbine engine according toclaim 1 wherein: the turbine is a first turbine, the compressor is afirst compressor, and the core shaft is a first core shaft; the enginecore further comprises a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor; and the second turbine, second compressor, and second coreshaft are arranged to rotate at a higher rotational speed than the firstcore shaft.
 9. A method of operating a gas turbine engine on anaircraft, the gas turbine engine comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox that receives aninput from the core shaft to drive the fan at a lower rotational speedthan the core shaft, the gearbox having a gear ratio of around 3.2 orhigher, for example between around 3.2 and around 3.8, wherein themethod comprises operating the gas turbine engine to provide propulsionunder cruise conditions such that a jet velocity ratio between a firstjet velocity exiting from a bypass duct of the engine and a second jetvelocity exiting from an exhaust nozzle of the engine core is within arange of from around 0.75 to around 1.0.
 10. The method of claim 9,wherein the jet velocity ratio is within a range from around 0.75 toaround 0.82 at cruise conditions.
 11. The method of claim 9 wherein thejet velocity ratio, R_(J), is defined as:$R_{J} = \frac{V_{B}C_{B}}{V_{C}C_{C}\eta_{LPT}\eta_{F}}$ whereV_(B) is the fully expanded first jet velocity, C_(B) is a thrustcoefficient of the bypass nozzle, V_(C) is the fully expanded second jetvelocity, C_(C) is a thrust coefficient of the core exhaust nozzle,η_(LPT) is an isentropic efficiency of a lowest pressure turbine of theengine core and η_(F) is an isentropic efficiency of compression of airinto the bypass duct by the fan.
 12. The method of claim 9 wherein thegearbox is an epicyclic gearbox comprising an input sun gear connectedto the core shaft, a plurality of planetary gears connected by a carrierarm and an outer annulus ring, the fan being connected to the carrierarm.
 13. The method of claim 9, wherein the fan has an outer diameter ofbetween around 220 cm and around 390 cm.
 14. The method of claim 13,wherein the fan has an outer diameter of between around 240 cm andaround 280 cm or between around 330 cm and around 380 cm.
 15. The methodof claim 9, wherein: the turbine is a first turbine, the compressor is afirst compressor, and the core shaft is a first core shaft; the enginecore further comprises a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor; and the second turbine, second compressor, and second coreshaft rotate at a higher rotational speed than the first core shaft. 16.The gas turbine engines of claim 1, wherein cruise conditions correspondto a forward Mach number of between around 0.7 and 0.9 at an altitude ofbetween 10000 m and 15000 m, optionally a forward Mach number of 0.85and an altitude of 10668 m.